Small gas turbine engine having enhanced fuel economy

ABSTRACT

A gas turbine engine ducted to have the air pass sequentially through a single stage axial-flow compressor, a single-stage hybrid axial-radial centrifugal compressor, a burner, an inward flow radial turbine, an axial-flow, high-pressure turbine, and then exit nozzles. The axial compressor and axial turbine are on a common core shaft. The centrifugal compressor and the radial turbine are on a common porous hub and form an integral, modular unit in which forces on the radial turbine blades drive the centrifugal compressor. The porous common hub permits air to transpire from the compressor to the turbine, serving to cool the turbine. Fuel is burned in annular, &#34;folded-comma&#34; shaped burners with multiple fuel-injection and ignition zones (two stage combustion).

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a divisional application of application Ser. No. 08/778,407,filed Jan. 2, 1997 (U.S. Pat. No. 5,832,715), which is a divisionalapplication of application Ser. No. 08/494,801, filed Jun. 26, 1995 nowabandoned, which is a divisional application of application Ser. No.08/100,297, filed Aug. 2, 1993, now U.S. Pat. No. 5,454,222, which is acontinuation of application Ser. No. 07/486,360, filed Feb. 28, 1990,now U.S. Pat. No. 5,253,472.

BACKGROUND OF THE INVENTION

1. Field of The Invention

My invention relates to the field of gas turbine engines and, inparticular, to smaller engines having enhanced fuel economy.

2. Prior Art

In comparison with reciprocating and rotary engines, gas turbine enginesoffer significant benefits in terms of small size and light weight. Interms of fuel economy, however, gas turbine engines have lagged behindthe other engines, particularly for the small engine sizes. Gas turbineengines have also been considerably more expensive, mainly due tointricate designs, close manufacturing tolerances, and the use of exoticmaterials. Data as to many of these prior art engines are found in "TheAircraft Gas Turbine Engine And Its Operation," United Technologies(Pratt and Whitney), 1988; and "Aircraft Gas Turbine Guide," (GeneralElectric (Aircraft Engine Group), 1980.

My gas turbine engines have a smaller size and lesser weight thancurrent turbine engines having the same power. Consequently, theyprovide less weight for the same power or more power for the same weightas prior engines. As a result, they offer greater range or flying powerand savings in fuel consumption, initial cost, and maintenance.

BRIEF SUMMARY OF THE INVENTION

The basic core structure of my engine is ducted to have the air passsequentially through a single stage axial-flow compressor, asingle-stage hybrid axial-radial centrifugal compressor, a burner, aninward flow radial turbine, an axial-flow, highpressure turbine, andthen exit nozzles. The axial compressor and axial turbine are on acommon core shaft. The centrifugal compressor and the radial turbine arenot separate units on a shaft, but form an integrated unit with a commonface.

The single-stage axial-flow compressor has a reduced setting angle forthe blades leading to a higher ratio than normal between the streamwisechord and the axial chord and a high compression tip velocity. Thisserves to optimize pressure rise and stage efficiency through sacrificeof axial mass flow. The pressure rise is multiplied with the followingcentrifugal compressor.

The first turbine in the series is a high-reaction inward-radial entry,axial-discharge transonic first stage driving a high-pressurecentrifugal compressor; it is followed by an axial flow high-impulsesecond stage driving the core shaft which is coupled to the low-pressureaxial compressor. The core shaft also provides shaft power to a gear boxfor power output or drives a variety of devices such as fans, boostercompressor, etc.

The combined centrifugal compressor and inwardly radial turbine have acommon face and so rotate together. The common face is formed of porousmaterial permitting air seepage from the compressor to the turbine,serving to cool the turbine. The seepage of air serves to remove stalledboundary layer air from the compressor and to cool the radial turbine,thus accomplishing a double function. Having a common face between thecompressor and turbine also eliminates windage losses for bothcomponents and hence improves their efficiency. This compressor-turbineunit is modular and, so, may be readily replaced as a unit.

Fuel is burned in annular, "folded-comma" shaped burners with multiplefuel-injection and ignition zones (two-stage combustion). The foldedcomma shape of the burners permits them to fit within and substantiallyfill the annular spaces radially outward of the turbines, providing fora more compact engine. It also results in the burners being surroundedon all sides by fluid ducts which use the heat; this is more efficientand eliminates the need to cool the engine cowl. The burners feed a setof nozzles for the high-pressure turbine. Due to the curved shape of theburners, infra-red radiation in the exit nozzles is substantiallyreduced.

The core engine is normally preceded by a grill-set, dust deflector andfilter, and is followed by a cross- and counterflow recuperator totransfer thermal energy from exhaust gases to compressed air flowing tothe burners.

A typical core engine of my invention would have a diameter of about0.35 m., and a length, including inlet ducts, recuperator, and exitnozzle of about 0.75 m. This engine would deliver about 600 kW (800 bhp)of power, with realistic component efficiencies, at about 54,000 rpm ofthe power shaft. The power may be used at either end, front or rear, ofthe engine.

My engine may be used for a turbojet engine, low-, medium-, orhigh-bypass-ratio turbofans, or in a turboshaft configuration. The basicdesign for the fans is the use of swept blades with long chords and highsolidity, avoiding the need for snubbers.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is a section of the core of my engine taken on a radius thereof;

FIG. 2 is a developed transverse section looking in a radial direction,showing my "folded-comma" burners;

FIG. 3 is a full transverse section, taken as indicated on line 3--3 ofFIG. 1, showing the centrifugal compressor 19 and its low-pressure bleedduct, together with the motor rotor and stator;

FIG. 4 is a full transverse section, taken as indicated on line 4--4 ofFIG. 1, showing the high-pressure bleed and the free vortex diffuser;

FIG. 5 is a full transverse section, taken as indicated on line 5--5 ofFIG. 1, showing the turbine rotor;

FIG. 6 is a full transverse section, taken as indicated on line 6--6 ofFIG. 1, showing recuperator ducting;

FIG. 7 is a perspective view, broken away to provide internal details,partially in longitudinal section, showing my core engine as part of alow-bypass-ratio turbofan; this Figure also shows an inlet grill, debrisblocker, and dust collector of a type usable with my core engine;

FIG. 8 is a section, taken on line 8--8 of FIG. 7, showing the axialturbine blades and the recuperator;

FIG. 9 is a section taken on line 9--9 of FIG. 7, showing an inlet guideblade; and

FIG. 10 is a longitudinal section of a medium by-pass ratio gearedturbo-fan with a two-stage recuperator.

DETAILED DESCRIPTION OF THE INVENTION

The primary elements of my core engine 15 are shown principally in FIGS.1 and 7. These include, in sequence, an axial flow compressor 17, acentrifugal compressor 19, recuperators 21 (not always used), annular"folded-comma" burners 23, a radial high pressure turbine 27, an axialflow turbine 29, and an exhaust nozzle 31. As stated, the centrifugalcompressor and radial turbine have on a common face formed of porousmaterial so that transpiration occurs between them, reducing thecompressor boundary layer and cooling the turbine.

In addition there may be the usual inlet grill 35, debris blocker 37,and dust deflector 39 (FIG. 7) and also a core shaft 43, and outputcoupling 45, and the usual bearings and air seals.

Axial Flow Compressor

The single-stage axial flow compressor 17 is a supersonic design withshocks in both the rotor 51 (with hub 53) and the stator 55 (with hub57). It has blades 41 with sharp leading edges, a ceramic portion 42(FIG. 9), and thin, controlled-diffusion airfoils.

Conventionally, axial-flow compressors have been designed to maximizemass-flow per unit frontal area, with some sacrifice in pressure riseper stage and/or stage efficiency. Because centrifugal compressors havelarge frontal area, axial compressor stages that are followed bycentrifugal compressors do not need to maximize axial mass flux density,but conventional engines have such axial booster stages scaled directlyfrom other axial flow engines.

Conventional wisdom has it that higher axial velocity allows greaterenergy to be added to the airflow for the same tangential velocity ofthe blades, and hence generates greater pressure rise across the stage.However, that viewpoint ignores the effects of blade angles that allowgreater energy addition and pressure rise for smaller axial velocities.Further, smaller axial velocities allow reduced inlet relativevelocities and Mach numbers, reducing inlet shock losses for transonicstages. Small axial velocities for high pressure rise do imply higherDiffusion Factors and DeHaller Numbers. However, small axial velocitiesallow longer blade chords that allow larger area ratios forseparation-limited diffusion angles, and hence allow larger diffusionfactors to be possible. Low axial velocity also prevents stalling of thefirst stage during start up.

In a conventional compressor during start up, at low rotational speeds,the pressure rise and hence increase in density across the first stagesof the compressor is rather small. Thus the low-density, i.e., highvolume, airflow reaches some velocity in the small passages in thelatter stages of the compressor that are designed to pass high-pressure,high-density flow. This sonic "choking" of the flow in the aft stagesreduces axial velocity in the forward stages, where the blades are setat angles appropriate for high flow. The blades in the forward stagesthus stall, suffering loss of performance and mechanical vibrations.

Damage to the blades during start-up in conventional engines isprevented by using high-power starter motors to quickly accelerate theengine through the stall during start up, although the problem remainswith many conventional engines. Use of a low axial velocity throughoutthe engine avoids the choking of flow in the latter stages of thecompressor. And, even if the flow was to choke, the reduced angle of theblades with the tangential direction can be less than the stalling angleof the blades, eliminating the problem of start-up stall.

My axial compressor stage 17 is designed to optimize a combination ofpressure rise and stage efficiency while sacrificing axial mass flow.This design also minimizes the vulnerability of the blades to stallbecause of the reduced setting angle of the blades at their leadingedges. I use, for example, a setting angle (between the blade camberline at the leading edge and the tangential direction) of about 15°; butthe angle can be lower. (The corresponding blade angle with the axialdirection is about 75°). Further, the blades can then have longerstreamwise chord for the same axial chord, for enhanced strength of theblades and/or reduced thickness/chord ratio for reduced shock losses.The ratio of streamwise to axial chord can range between about 4 toabout 1.

One common problem with axial compressors is the rapid growth ofstagnant boundary layers on the hub surface of the compressor rotor dueto the centrifugal field created by the rotor. This causes stall of theblades in the hub region and reduction in the compressor efficiency. Myengine alleviates this effect by having a sharply-curved flow passage 55in the compressor so that the passage appears concave radiallyoutwardly. This curvature, as shown, will impose a centrifugal flowfield due to the flow velocity, inducing a tendency for the air streamto adhere to the hub contour. This helps counteract the radially outwardflow field caused by the rotor.

Suppression of the boundary layer growth near the hub surface is alsoprovided by allowing air to leak through the gap 54 between the rotorhub 53 and the stator hub 57. This leakage air reduces the boundarylayer thickness near the hub surface and also provides the air forcooling of bearings and motors/generators 67 and 68 on the shafts.

The wall boundary layer near the axial shroud 63 is tapped off into alow-pressure bleed duct 65. This low-pressure bleed air may be used forcooling fuel pumps, avionics, etc. on the engine or user vehicle or forpneumatic flow control on vehicle surfaces.

An additional bleed gap 79, just before compressor blades 77, providesbleed air for cooling the bearings.

My axial compressor design provides for a stage pressure ratio of about2:1, with a stage efficiency of about 90% (compared to the usualefficiency of about 85%). It achieves this by having a tip velocity ofabout 450 m/sec (1500 ft/sec) with a rotor speed of about 54,000 rpm.Stage efficiency is maximized through the use of negative pre-swirl,swept leading edges, flutter-resistant low-aspect-ratio blades, filletsat blade roots, relatively high hub/tip ratios, radial equilibrium aidedby duct curvature and by, low axial velocity, as noted above.

The shroud 63 of the axial compressor has a porous surface facing theair flow path. The pores will trap dust particles in the air that impactthe shroud due to the rotational flow field generated by the axialcompressor rotor. The dust will thus be prevented from blocking thetranspiration passages in the centrifugal compressor and turbine system.

Further, the stator section following the axial compressor has dustcollecting airfoils that have porous surfaces on the pressure side.These porous surfaces will trap dust particles impacting the same due tothe camber of the flow path.

Centrifugal Compressor

The centrifugal compressor 19 is a hybrid axial-radial design withtransonic inlet velocity. The blades have sharp leading edges at theinlet and backward curvature at the exit for high efficiency and flowstability. Compressor 19 and radial turbine 27 (described below) have acommon hub 75 which is concentric with but separate from core shaft 43.

The hub 75 carries impeller blades 77 on surface 26 and the turbineblades on surface 146. Hub 75 is porous, allowing boundary layer air onthe hub surface to seep through the hub material into the radial turbine27. This porous material can be a ceramic like silicon nitride or,preferably, a fiber-reinforced ceramic like silicon carbide whiskers insilicon nitride with a high fiber content and low matrix content. Theporosity should be sufficient to suppress compressor boundary layersand, to the extent of the transpiration, to cool the turbine.

This air seepage suppresses growth of the boundary layer on the impellerhub 75, prevents heating of the impeller 77 by the turbine 27 (whichwould otherwise reduce compression efficiency), and provides for coolingof the turbine to retain turbine strength despite the high temperatureof the gases in the turbine flow passages.

Ideally, the porosity of hub 75 will be adequate to providetranspiration that will reduce the boundary layer to approximately zero.Less porosity will provide less transpiration causing the compressor tobe less efficient due to the presence of some remaining boundary layer;more porosity than that will result in too much transpiration, wastingair and causing loss of mechanical power.

This integral compressor-turbine unit is modular in that after wear, orif the pores of the hub become clogged with dust despite the dustseparators, discussed below, the entire compressor-turbine unit may besimply replaced as a whole.

The centrifugal compressor 19 has a high pressure bleed duct 59 on thecentrifugal shroud surface 64 after the impeller. This bleed duct tapsoff any boundary layer on the shroud surface of the impeller and alsocan provide air for pneumatic flow control on the user vehicle, orpressurization of the main cabin, and the like.

Centrifugal compressor 19 has a design pressure ratio of about 5:1 and atarget stage efficiency of about 80%. The design tip speed is relativelyhigh, about 600 m/sec (2,000 ft/sec) at a rotational speed of about60,000 rpm.

These high rotational and tip speeds are possible, even with thisbackward-curved design, because of low radial and tangential stresses.The radial stresses are low because the impeller 77 is made oflow-density fiber reinforced ceramic material, such as silicon carbidewhiskers in silicon nitride, having a density in the order of aboutthree grams per cubic centimeter, and because of the porous nature ofthe common compressor-turbine wheel 87 (FIG. 8). This results in lowcentrifugal stresses. The tangential stresses are low because thecompressor is not shaft driven, as in conventional compressors, but isdriven by facial contact of the impeller vanes 77 along their lengthwith hub 75, for direct transfer of forces between the turbine 27 andthe compressor 19.

The stator section 81 of the centrifugal compressor has exit guide vanes81 following the principal flow direction of the air to act as adiffuser and convert the high axial-radial-tangential velocity of theair leaving the impeller into lower axial velocity and higher pressure.The diffuser inner shroud 85 and the vanes 81 also have a porousconstruction, normally being made of the same material as thehub/interface 75, to reduce diffuser boundary layers, reduce heattransfer from the turbine nozzles, 30 and allow transpiration cooling ofthe nozzles.

The Recuperator

There are instances, such as with turbo-shaft engines or ahigh-by-pass-ratio turbofan, where it is advisable to have arecuperator. In a recuperator a significant amount of thermal energy istransferred from the hot exhaust gases leaving the axial flow turbine 29(described below) to the high pressure air of the compressor 19 beforethe air is ducted to the burner 23.

The high pressure air from compressor 19 passes through stator 81 and iscarried by annular duct 93 to recuperator 21 (FIG. 7) located near therear of the engine 15. (The recuperator is not shown in FIG. 1, butwould be positioned just aft of the core engine). The recuperator itselfis a cross-flow shell and tube design and the hot gases flow in aquasi-axial direction from the axial flow turbine 29 to the exhaustnozzle 31. Air from the compressor 19 is ducted radially inwardlythrough a large number of narrow tubes 95 from the outer annular region97 to a chamber 99, near to and concentric with the axis of the engine,and then radially outwardly through more narrow tubes 95 to anotherouter annular region 103. This radial-in and radial-out set of tubes 95,along with the associated ducting, forms one unit of the recuperator.

Many engines, including the core engine, could have one or more suchrecuperator units. For example, the high-pressure, low-bypass-ratioturbofan of FIG. 7 has only one such unit, while a moderate-pressure,medium-bypass-ratio turbofan as shown in FIG. 10 might contain two; anda turbojet none. This modular usage of one or more recuperator unitsoptimizes the combination of overall engine pressure, ratio andrecuperator effectiveness for optimum engine efficiency and weight.

FIG. 10 shows a second recuperator stage 22 added after recuperatorstage 21. Duct 97 has been modified to feed recuperator stage 21 with aninterstage duct 104 added to transfer flow from recuperator stage 22 torecuperator stage 21. Chamber 100 performs the same function for stage22 as chamber 99 does for stage 21.

The recuperator provides significant gains in fuel economy by making useof thermal energy in the exhaust gases to preheat air before combustion.It also reduces the infra-red signature of the engine by, first,reducing the exhaust gas temperature (for a fourth-power reduction ininfra-red emission), second, blocking a direct line of sight from thehot turbines to the exhaust nozzle, and, third, by having the hotturbine section surrounded by the two annular ducts 93 and 103 runningto and from the recuperator (FIG. 7). These ducts also recover some ofthe thermal energy that would otherwise be lost through the casing 30 ofturbine 29. By reducing heat that needs to be added in the burner, therecuperator also reduces the size needed for the burner.

The "Folded-Comma" Burner

I refer to my burners as "folded-comma" burners because these burners 23have the appearance of a folded comma as viewed in FIGS. 1 and 7.

The folded comma burner unit 23, is located circumferentially of thenook of the radial turbine 27. A typical engine 15 may have six burnerunits spaced annularly about turbine 27. Each burner 23 includes twodistinct types of burning regions: one, a pilot region formed by awide-angled, high-turbulence diffuser 113 (shaped as an open-endeddiverging two-dimensional duct) in the feeder duct 111 and, two, a mainregion having a pair of bulbs 115 (shaped like the main body of acomma). The two main regions in each burner receive air through thespace between the inside surface of feeder duct 111 and the outsidesurface of diffuser 113. Each bulb-shaped main region of the burnerreceives air from two ducts 111, just as each duct 111 feeds two bulbs115. The pilot region includes fuel spray nozzle 121; and the mainregions, fuel spray nozzles 123. The nozzles are fed fuel through fuellines 127 leading from fuel pumps 125 (located radially outwardly ofaxial flow compressor 17).

Nozzles 121 and 123 are preferably ceramic, possibly high densitysilicon nitride. This allows the fuel nozzles, which have low strengthrequirements and are internally cooled by flow of fuel, to operatesubmerged within the burning gases. The nozzles inject a mixture of fueland air, and are purged by air after fuel shut-off to prevent cooking offuel in the nozzles. Alternatively, ultrasonic pulses in the fuel linemay be used to create atomizing pressure pulses as well as keep the fuelinjectors free from coke and other fuel residues.

The burners should best be made of woven carbon-carbon composites, andpossibly have several layers of ceramic coating. Carbon-carbon exhibitshigh strength to temperatures higher than adiabatic flame temperatures(about 2260° K./3600° F.) for typical fuels, provided it can beprotected from oxidation. One possible material would be wovencarbon-carbon with SiC conversion coating and a coating of ZrO₂,possibly mixed with HfC and ZrB₂. Other coating materials include ZrC,Ir, and SiO₂.

In operation of the burners, hot, compressed air from the recuperator 21is ducted forward through feeder duct 111 to the burners 23, and the tworegions 113 and 115 are sprayed with fuel through nozzles 121 and 123.The hot products of combustion then pass to the rotor 141 of radial flowturbine 27 by way ducts 131 (the folded tails of the "commas") andnozzles 139. Turbine 27 includes rotor blades carried in throat 143 byhub 75. The mean radius of throat 143 is larger than the mean flowradius halfway through the turbine.

As can be seen, the basic design philosophy behind my burner design istwo-stage combustion with a pilot region and a main region. The pilotregion burns part of the fuel under fuel-rich conditions, with theresulting oxygen deficiency minimizing the production of oxides ofnitrogen.

Flame stability is provided by the turbulence in the rapidly divergentduct causing recirculation of burning gases. The main region would thenhave the fuel injected into the hot burning gases flowing from the pilotregion and so would achieve rapid combustion with very little residencetime in the burner. Any soot formed in the pilot region would beoxidized in the high temperature main region; at the same time,residence time in the main region would be small enough to minimizeformation of NOx.

The flow path for gases is designed to utilize heat transfer from hotgases to compressed air and enhance efficiency. Air leaving centrifugalcompressor 19 in duct 93 is preheated by the air entering radial turbine27 by duct 131 since ducts 93 and 131 have a common wall. The air induct 93 is further preheated because the duct then has a common wallwith air in duct 111 leading from the recuperator. It should be notedthat the ducts with the coolest air are on the outside of the engine,reducing heat loss. This provides constructive shrouding and reduces theneed for cooling air for the cowl, resulting in a self-cooling cowl.

Also, as can be seen, my design utilizes annular space radially outsidethe turbines, which space has not previously been used. The folded-commashape of the burner reduces the cross-section, so the engine is smaller.Having the burner surrounded on all sides by fluids which use the heatmeans that my engine requires less heat to develop the necessary energy,and so it is more efficient.

The Radial High-Pressure Turbine

As stated above, the high-pressure centrifugal compressor 19 is mountedback-to-back and on the same hub 75 as the inward flow radial turbine27. As can be seen, its blades 141 are mounted on surface 146 of hub 75.Thus, rotation of turbine 27 serves to rotate compressor 19, and theyrotate at a common speed.

Turbine 27 is fed by a ring of nozzles 139 which receive hot combustiongases from duct 131; and the nozzles convert the annular axial flow ofthose high pressure gases into a substantially tangential andinwardly-radial high-velocity flow that impinges on the inward flowradial turbine rotor 141. The nozzles share the outer shroud 85 incommon with the diffuser inner shroud 85, the shroud being made of aporous material to allow transpiration of boundary layers from thediffuser to cool the nozzles.

The gases leave the nozzles at a supersonic absolute velocity of aboutMach 1.3, but, because of the high rotational speed of the rotor 141(about 60,000 rpm), have a subsonic velocity of about Mach 0.7 relativeto the rotor. The gases expand further in turbine rotor 141, through thesonic throat region 143, into larger discharge area 147, and leave therotor, at supersonic speed, in a direction reversed almost 180° from theentrance direction. Turbine rotor 141 is therefore driven by the impulsecaused by the change of direction of the relative velocity as well asthe reaction caused by the high exit relative velocity of the gaseseffluxing from the turbine.

Turbine 27 is designed to produce only sufficient power to drive thecentrifugal compressor 19 plus enough power to drive a high speedelectric generator 67, 68 integral with the rotor 141.

The compressor-turbine hub 75 is made of ceramic-fiber reinforcedceramic-matrix composite (e.g., SiC whiskers in Si₃ N₄ matrix) integral,as stated above. This compressor-turbine hub has a high fiber contentand low matrix content which gives it high tensile strength to withstandrotational stresses and relatively high porosity to allow seepage ofairflow from the compressor to the turbine. This seepage flow fromsurface 76 to surface 146, which is at a rate of about 2% of engine airflow, suppresses boundary layers in the compressor, reduces heating ofthe compressor by the turbine, and provides for transpiration cooling ofthe turbine so that it retains adequate strength at its high operatingtemperatures.

The Axial-Flow Turbine

The high-velocity discharge from the high-reaction radial turbine 27,leaving exit region 147, drives single-stage axial flow turbine 29. Thegases leaving axial-flow turbine 29 are of low tangential velocity andhave a moderate axial velocity of about 300 m./sec. (Mach 0.5)

The axial-flow turbine 29 has blades 151 and drives the power outputshaft 43 of the engine through hub 153. This output shaft is concentricwith and central to hub 75 the hollow-shaft centrifugal compressor 19and the radial turbine 27. Hub 75 and shaft 44, however, rotateindependently of one another. Power output from the axial-flow turbinecan be used at either the front or the rear of the engine, with orwithout gear boxes.

Hot gases from the axial-flow turbine 29 pass through the recuperator 21(FIG. 7) where, as described above, heat is transferred from the effluxgases to the compressed air flowing from the compressor 19 to theburners 23. The hot efflux gases finally exit from the engine through aconverging nozzle 31 (FIG. 7) at the rear of the engine. Alternatively,the nozzle may be converging-diverging for high-pressure turbojet-typeengines.

Bleed Ducts: Active Through-Flow Control

Engine 15 has bleed ducts 65 and 59 for low pressure air from the axialflow compressor 17 and high-pressure air from the centrifugal compressor19, respectively. Bleed duct 54 between the rotor hub 53 and stator hub57 allows boundary layer air to cool bearings and motors/generators 67and 68. For a typical engine application, the high-pressure bleed airmay be used for cabin heating and pressurization, and the low pressureair may be used for blowing on airfoils (for augmented lift) on the uservehicle. In addition, the bleed system provides active throughflowcontrol for the compressors to enhance their performance, efficiency,and stability, as noted below.

During starting or operating at low speeds, the first stages on typicalengines generate reduced compression and reduced gain in density, whichcauses the latter stages to choke out, which in turn causes the firststages to stall. My engine prevents stalling the axial compressor at lowspeeds by having the low pressure bleed valve open to maintain highvolume flow rate through the compressor. Similarly, if the centrifugalcompressor has a tendency to stall because of choking of the exit guidevanes, the high-pressure bleed valve may be partially opened to reducevolume flow through the exit guide valves.

Inlet Area

FIG. 7 shows my core engine being used in a low-bypass-ratio turbofan.Inlet grill 35, the "bird blocker" grill, has a large mesh, about 5 cm(2 inches) grid size. Debris exits through slots 36. Grill 35 isfollowed by "insect blocker" grill 37 of smaller mesh, about 1 cm (0.4inch). The air then passes through a low bypass ratio fan 38, driven byshaft 43, a portion of it entering my core engine 15 and the remainderleaving by fan-jet nozzle 40.

The inlet guide vanes for the first compressor have porous surfaces onthe pressure side of the vanes. These porous surfaces will trap dustparticles that impact the pressure surfaces of the guide vanes due tocamber of the air flow path created by the vanes. The inlet guide vaneswill thus form an effective dust filter with a very low pressure drop,and will be replaceable as a module. The porous surface here, as withthat of shroud 63, may, for example, be made of a felt-type mattedsurface or open-cell plastic foam.

I claim:
 1. In a gas turbine engine having a compressor, a burner, and aturbine section, wherein the improvement comprises:the turbine sectionhaving a turbine rotor with an inward substantially radial entry and anoutward substantially axial discharge, wherein the turbine rotor has agas flow path with a middle section, about halfway along the length ofthe flow path through the turbine rotor, that has a mean radius ofcross-section of the flow path, the mean radius being measured from anaxis of rotation of the turbine rotor, wherein the mean radius of thegas flow path at the middle section is smaller than the mean radius ofthe gas flow path at the entry to the rotor and is smaller than the meanradius of the gas flow path at the discharge from the rotor.
 2. Anengine as in claim 1 wherein the middle section is located closer to theaxis of rotation of the turbine rotor than the entry.
 3. An engine as inclaim 2 wherein the middle section is located closer to the axis ofrotation of turbine rotor than the discharge from the rotor.
 4. Anengine as in claim 1 wherein the turbine rotor is suitably sized andshaped to convert gases traveling into the entry at a subsonic velocityrelative to the turbine rotor into a supersonic velocity relative to theturbine rotor as the gases exit the discharge.
 5. In a gas turbineengine having a compressor, a burner, and a turbine section, wherein theimprovement comprises:the turbine section having a turbine rotor with aninward substantially radial entry, an outward substantially axialdischarge and a middle section therebetween, wherein the turbine rotorhas a gas flow path between the entry and the discharge with an innerwall, an outer wall, and turbine blades therebetween, the inner andouter walls of the gas flow path forming a cross-sectional area with ageneral annulus shape, wherein the cross-sectional annulus area of thegas flow path at the middle section is smaller than the cross-sectionalannulus area of the flow path at the discharge from the rotor.
 6. Anengine as in claim 5 wherein the middle section is located closer to anaxis of rotation of the turbine rotor than the entry.
 7. An engine as inclaim 6 wherein the middle section is located closer to the center ofrotation of the turbine than the discharge.
 8. An engine as in claim 5wherein the turbine rotor is suitably sized and shaped to convert gasestraveling into the entry at a subsonic velocity relative to the turbinerotor into a supersonic velocity relative to the turbine rotor as thegases exit the discharge from the rotor.